N88-27162 - NASA

N88-27162

AEROELASTIC CHARACTERISTICS OF TIIE AH-64

BEARINGLESS TAIL ROTOR

D. Banerjee

Chief, Aeromechanics R&D

Hughes Helicopters, Inc.

Culver City, CA 90230

Abstract

A Composite Flexbeam Tail Rotor (CFTR)

with a s}ructurally and aeroelastically unique hub

desit_n has been developed at Hughes Helicopters,

Inc. (HHI) for the AH-64, Advanced Attack Heli-

copter. The full scale rotor has been success-

fully tested in the wind tunnel over the full steady

sideslip envelope of the AH-64. The test program

has defined the performance, loads, and dynamic

characteristics of the CFTR for rotor speeds up

to I. 0 N R and airspeeds up to 197 knots. Unique-

ness of the design is reflected in its patented hub

design. The elastomeric shear attachment of the

flexbeam to the hub results in a soft-inplane

S-mode and a stiff-inplane C-mode configuration.

The properties of the elastomer have been

chosen for proper frequency placement and stable

damping of the inplane S-mode. Both frequencies

are well separated from the l-flap frequency.

The stress-critical pitch case/blade interface has

been carefully designed to minimize loads. The

flexbeam spanwise thickness and Width distribution

have been tailored for near-uniform corner

stresses. The I/rev chordwise load is main-

tained within the flexbeam and is not transferred

to the hub. The Z/rev chordwise loads are trans-

ferred to the hub after significant attenuation due

to hub shear pad damping and separation of the

reactionless l-chord frequency from Z/rev. The

carry-through design of the flexbeam across the

rotor hub allows the flexbeam to deform within

the hub to reduce the hub loads to a minimum.

Kinematic pitch-lag coupling is introduced to

improve the first cyclic inplane C-mode damping

at high collective pitch.

Presented at the Integrated Technology Rotor

(ITR) Methodology Workshop, NASA Ames

Research Center, Moffett Field, CA,

June 20-Zl, 1983.

1.0 Introduction

Hughes Helicopters, Inc. (tIHI) has designed,

fabricated and successfully wind tunnel tested a

Composite Flexbeam Tail Rotor (CFTR for the

AH-64 Advanced Attack Helicopter.

Over the past several years, a varlety of

bearingless tail rotors have been developed. The

CFTR is a bearingless rotor whose design

features have benefited from recent advances in

composites technology and lessons learned from

research into the basic characteristics of bear-

ingless rotors that have to be addressed to

achieve a successful design. Reference I

describes the experimental development of a

bearingless rotor and shows that a rotor system

whose coupling effects are not _vell understood

can run into fundamental dynamic instability

problems Instabilities encountered in the design

were:

I) Inplane C-mode instability.

2) Inplane S-mode instability.

3) Stall flutter in the third flexible mode

(torsion).

4) Stall flutter in the fourth flexible mode

(second flap).

This reference also provides valuable infor-

mation on the effect of key parameters such as

blade sweep, tip weight, kinematic pitch-flap

coupling, flexbeam width, etc. , on the dynamic

and aeroelastic behavior of the rotor. The choice

of flexbeam geometry was found to be crucial to

the level of flexbeam loads, and hence, the per-

missible amount of the kinematic pitch-flap

coupling, which influences the flexbeam fatigue

loads. In Reference 2, a hingeless rotor had

carefully designed flexbeam and was inherently

stable. A closer look at this concept raised

several questions regarding the "optimality" of

the load path in the rotor. In Reference 3, the

RT0 IGFNAL PAGE IS

OF POOR QUALIT_

279

rotorsystemcnco,mtcr,'d;tr, i_slahility involvin_

the first flap/ch,H'd _,,M,' al ,,_ud,,ralc (,dh'_tiw:

pitch. The rotor d,.scrih,.,I i,, R,.I,.r,..,," 4

encountered flap-la_ fr,'qu,'m y ,,>.l,.nc,_u_c aud

resultant instability whicl_ wan ,.li_,_iual,,d by

ct_anging the pitti_-tlap _ _mplir_lZ (6?) Ir,n,_ ;J _on-

vcntional value ol -$5 t_ -4q d_._r,.'s (flap up

induces pitch down), to t _5 dc_r_.,'s, thus reducing

the first flap frt_qm'ncy t_) bclov, I /rcv. IIuwcw'r

care had to be excrcised in th_ use of such pitch/

flap coupling sin(t it can lt'ad to static diw'r_(:,:

in flap/pitch. Tlu' rotor loads and perforH_anc,,

characteristics resulting from the varyin_ 6_

were not addrc_sscd.

These rotors can be generally categorized as

stiff-inplane or s0lt-inplane rotors. Typical

problen_s associated with stiff-inplane rotors

have })coin:

1) Inadequate structural stiffness in the

flexbeam to ensure adequate separation of

1-chord and 1-flap frequencies. This generally

results in coupled flap-la_ instability (Refer-

enc,_ _ 1 ).

2) Since the hub and drive system torsional

stiffness lower the frequency of the 1-chord

reactionless and collective modes, they have to

be taken into account in sizing the flexbeam

chordwise stiffness characteristics to avoid

coalescence of the 1-chord and 1-flap modes.

3) In ensuring good separation of the 1-chord

and l-flap modes, the 1-chord frequency is gener-

ally laced high (between 1. 5/rev and 1.7/rev).

Dynamic amplification of 1 /rev and 2/rev Coriolis

bending moments result in high 1 /rev and Z/rev

chordaise fatigue loads in the flexbeam.

4) In order to accommodate the high loads

of a stiff inplane rotor, a relatively stiffer flex-

beam is required. This also increases the

torsional stiffness of the flexbeam resulting in

Sigher lorsional loads on the control system.

Soft-inl)lane rotors have potential problems of:

l) Dynamic coupling of the rotor anti sup-

port structure resulting in "ground resonance"

type problems.

2) Structural loads in the flexbeam of a

bearingless rotor could determine a lower bound

on the flexbeam stiffness, and hence, the l-chord

frequency of the rotor blade.

With the above concerns in mind, the

Con_t)osite KlexhL.am Tail Rotor (CF-TR) has been

dew'loped at ttughes Ilelicol)ters, Inc. It has a

structurally tailored flexbeam chordwise stiffness

distribution to locate the cyclic 1-chord frequency

above l/rev, and the flexbeam is mounted to the

hub between elastomcric "soft" supports whose

stiffness and damping are tailored to locate the

collective and reactionless 1-chord frequencies

below 1/rev. A description of the rotor design

and dynamic characteristics are. presented in

Sections g and 3, respectiwdy.

g.0 CFTR - Description

An exploded view of the CFTR is shown in

F'ig. 1. This shows that tiae axes of the blade-

pair assembly arc perpendicular to each other,

and arc separated axially so one flexbeam may

cross over the other. Tile CKTR has upper and

lov, c_r hub plates whicln sandwich the blade-pair

assembly. The hub assembly is bolted to the tail

rotor drive shaft. The flexbeam extends from the

tip of one blade, across the hub, to the tip of the

opposite blade. Bending and twistin_ motion of

the flexbeam, betv_ecn the edge of the hub and the

inboard end of the blade, provides the fundamental

flap, lag, and torsional motions of the rotor

blades. The flexbeams are attached to the hub

plates through elastomeric shear (inplane) pads.

The laminated elastomeric pitcln shear support

aligns the pitch case with respect to the fiexbeam.

The pitch horn is bolted to the trailing edge of the

pitch case. The Sl)anwise location of the pitch

link attachment is adjusted for an effective pitch-

flap coupling (83 ) of -35 de}trees (pitch down with

flap up). The pitch link is inclined to provide

negative pitch-lag coupling (64 positive: pitch up

with blade lag) to augment inplane dampin_ at l_igh

collectiw' pitch and rotor speed. A brief descrip-

tion of each component follo_s.

28O

.Jr /

,, ,L

Fig. 1 CFTR assembly

OKIGiz _,_ _ .....:-"--,-,!.. IS

ON POOR QUALITY

2. l Flexbeam

OF POOR QUALIT_

The heart of the CFTR is the fiberglass/

epoxy flexbeam that carries across the full span

of each blade-pair assembly and attaches the two

blade sections of each blade-pair assen_bly to

each other and to the hub. The flexbeam, which

is of rectangular cross-section is built of layers

of S-glass/epoxy with the filaments oriented

+5 degrees to the spanwise axis. S-glass was

selected for its good fatigue strength, relatively

high elongation, and low modulus of elasticity.

Fiber orientation of 4-5 degrees was selected as

having a good fatigue strength and low torsional

stiffness combined with the inplane shear strength

to carry the driving torque and inplane blade

loads. The spanwise distribution of flexbeam

width and thickness is configured for near uni-

form spanwise distribution of combined corner

stresses while maintaining a low structural

torsional stiffness.

The flexbeam is formed as a flat beam that

operates in the untwisted condition when the blade

is producing design lift at 03/4 = 8 degrees so

that the torsional stress within the flexbeam is

minimized.

2.2 Hub

The hub, as shown in Figs. 1 and 2, consists

of upper and lower hub plates which sandwich the

flexbeams between elastomeric inplane shear pads.

Each set of pads is clamped between two load

carrying beamlike structures; an upper hub plate

"cross beam" and the "cross beam" stiffener of the

lower hub plate. These beams carry shear loads

due to preloading and reaction loading of the pads

to support points on their ends. The pads them-

selves consist of an elastomeric section bonded

to a thin aluminum plate which in turn is bonded

to the flexbeam. Four anchor bolts (two on each

end of each shear pad) attach the pads to the lower

•

,ROSS 8jAM" _-,H[AR _ANI:L

BRACED STIffENERS

Fig. 2 Hub design

hub plate which carries all the reaction loads to

the drive shaft. The elastomeric pads provide a

soft mount between the flexbeam and hub and are

designed to allow the flexbeam to bend with

respect to the rigid hub and to keep the primarybending moments within the flexbeam _here the

filaments art" oriented to accommodate them. In

addition, the hub, which is of hollow construction,

is designed to minimize the load path. These

features art, shown schematically in Fig. 3.

• FLA PWISE

TAI:_R[O EL[XBE_M CONTROLS BENDING STRESSES

EL ASIOMER CONTROLS FL[XBE/LM TO-HUIB LOADING

• bREV CORIOLIS-CHORDWISE "C" MODE

ELASIOMER ALLOWS FLE×BEAM BENDING

LOADS REMAIN IN FLEKBEAM MINIMAL TRANSFER TO HUB

• ?JREV CORIDEIS CHORDWISE "S" MODE

INTER BtADE-PAIR LOADS SHORT I OAD PATH

- ELASTDP'AER DAMPS SCISSORS _AOT/ON

Fig. 3 Hub design criteria

In the flapwise direction, the flexbeam is

designed for transfer of minimal bending moment

loads into the hub as a result of the flexbeam

taper and bending within the hub. The elastomer

is clamped to preload it and ensure that it always

has a net compression load. All flap bending

loads are transferred between the flexbeam and

hub through compression in the elastomer. The

loads are transmitt_d by the upper hub plate

"cross beam" and the lower hub plate "cross

beam" stiffeners to the shear panel braced stiff-

eners (Fig. l). These stiffeners are very deep

and, therefore, are structurally very efficient for

carrying the loads. The bolts for attaching the

shaft flange to ti_e lo_cr hub plate are anchored

at the intersection of these stiffeners with the

central pocket. This results in the shortest

possible load path.

Three predominant chordwise loads are

encountered. The first is the steady driving

torque which is reacted by the elastomer in

shear. The other two result from Coriolis forces.

The hollow hub allo_s the 1/rev Coriolis bending

moment loads to be carried in the flexbeam

instead of being transferred into the hub. The

Z-rev Coriolis moments result in the inplane

scissors S-type motion in _hich the adjacent

blades work against each other as shown in the

2S1

lower sketch of Fig. 3. In this case, tile loads are

taken in sh_,ar through the" elastomers and through

short load paths across the rugged corners of the

hub.

2. 3 Pitch Case

The pitch case is a _et-filament wound fiber-

_lass epoxy hollow structure that fits around, and

is bonded to the flexbeam and blade root _}Lerc

these three components intersect. Inboard of the

blade root, the pitch case enlarges to give the

flexbeanT room in which to twist as the blade

feathers (Yig:. 4). The pitch case tapers in the

spanwise direction (Fig. 4) to reduce the flapwise

stiffness (without sacrificing torsional rigidity).

This mini,uize_ the bendiu,_ u_O,nent in the pitch

case/blade root attachment induced by the pitch

Shear support anc1, hence, the resultant bending

stresses. Near the inboard end of the pitch case,

a hoop-wound stiffening ring provides the strength

required to support the pitch horn and the elasto-

n/eric shear support loads.

PITCH

SHEAR

/ SUPPORT

PITCH HORN z (SNUBBERI

_13 : - 350

EFFECTIVE FLAPPING HINGE

FOR CONTROL GEOMETRY

ADE

PSTCH CASE

ELASTOMERIC HUB PITCH SHEAR BLADE

SHEAR PADS SUPPORT ROOT CAP

HUB- __/ / (SNUBBER)]_, _I,-7*r11_

...... Y;LE×_EA2''_BLADE"" _ M N MAL R3"NTER NG MINIMAL PITCH SHEAR SUPPO

EFFECTIVE FLAPPING HINGE CASE REQUIRED INDUCED BENDING MOMENT INFOR CONTROL GEOMETRY PITCH CASE BLADE ROOT

Fig. 4 CKTR blade root ,_eometry

2.4 Pitch Shear Support ("Snubber"}

The elastomeric pitch shear support is a

laminated n_etal/elastomer device that is stiff

N_ith respect to radial loading, bEEt soft in torsion

and inplanc shear. It centers the pitch case with

respect to the flGxbeanl. Its spanwise location is

kept _ell outboard, beyond the region of maximum

flap bending curvature in the flcxbcam. This

n_inimizes the rotal ionill deflection of the pitch

case relative to the flcxbean_ as seen in the 1owE'E"

vietN of Fi_. 4, and so ininindzcs pitch shear

SUplx) rt-indu('t'd ben(ling _non_ents ;it [he pL)int

_here ti_e pitch case, flcxbcam, and blade join at

the bladp root station.

Z. 5 Blade

The primary material for the wet filament

wound blade structure is Kevlar-49/epoxy.

Unidirectional fibers with maximum tensile

strength and modulus are used for leading edge

obstacle strike protection, and for the trailing

edge longo that carries high axial loads and has

high stiffness. The airfoil-shaped blade section

is a multi-tubular Kevlar/epoxy structure that is

bonded around the flexbeam (Fig. 5). A C-shaped

channel is added in the aft airfoil region to stiffen

the outer skin. The leadin_ edge balance weight

is a multiple-rod mohled construction. The small

diameter rods easily conform to twisted contour

of the leading edge. The portion of the leading

edge cavity between the leading edge balance

weight and Kevlar spar tubes is filled with

syntactic foam.

POLYURETHANE

STAINLESS STEEL KEVLAR!EPOXY SKIN AND

EROSION STRIP //ALU_IINUM LIGHTNING SCREEN

FIBERGLASS EPOXY

,' FLEX BEAM ,"'?_,/_>¢. , ,"VEAR'EPOX,TRA,E,NGEOOEsGCASSEP0X_"---'_'_ "S

ELECrROIVERMAL

DEICER KEVLAa/EP ,

SPAR TUBES KEVLARIEPOXY

"C' CHANNEL

Fig. 5 CKTR blade cross-section

The blade has a -9 degree twist, and is

positioned about the flexbeam so that when thE,

flexbeam is untwisted, the blade pitch angle at

3/4-radius is 8 degrees. The orientation of the

flexbeam with respect to the blade chord at differ-

ent radial stations is shown in Fig. 6.

_. 0 CYTR - Dynamics

The fundanwntal mode of instability for bear-

ingless rotors has been shov, n both analytically

and experin'.entally to be associated _Kitln the

couplin_ betv, etm the first flap and the first

inplane (reactionless and cyclic)modes (Refer-

enccs 1, _, 4, 5, (, and 7). For bearinRless tail

rotor designs (l{cferences 1, Z and 4), the inplane

frequency generally lies between 1 and 2/rev,

with the rcactionless (S} mode frequency slightly

282

ORIGINAL "_ _P rC

2OOR qU.:.,_,, : :,.

<,.'. _3F,_,,_ Q O.::_LiTy.

RADIAL STATION (r/R)0.13 (EDGE OF HUB)

/_1_ o

0.20

0.27

0.32 t5 ° 5.70

Fig. 6 Blade/pitch case/flexbeam

cross -sections

lower than the cyclic (C) mode frequency - the

difference depending on the hub configuration and

the rotor pylon structural properties. Both an

increase in collective pitch and (conventional)

negative pitch flap coupling tend to bring the first

flap and the first inplane frequencies closer

together, by increasing the first flap frequency

and lowering the first inplane frequency. This

often results in the lightly damped first inplane

modes (both the reactionless and cyclic) becoming

unstable. Typical solutions to the above problem

have been the stiffening of the flexbeam in the

chordwise direction {Reference 1) and the use of

positive pitch flap coupling (Reh'rences 4 and 6)

to separate the modes. These solutions have been

applied with limited success because first,

structural design considerations put a limit on the

chordwise stiffness of the flexbeam, and second,

even though a stable rotor system was presented

in Reference 4 (with positive pitch-flap coupling),

similar experimental effort in Reference l showed

the presence of a stall-induced flap-lag-torsion

large amplitude limit cycle instability.

At HHI, the above dynamic problems have

been solved for the CFTR by lowering the S-mode

inplane frequency below 1/rev (soft inplane) while

maintaining the C-mode inplane frequency above

1/rev (stiff inplane} and well separated from the

first flap frequency. Some of the design param-

eters that resulted in this dynamically unique

bearingless tail rotor design are discussed below.

3. 1 Flexbeam to Hub Support

By supporting the flexbeam to the hub

through elastomeric hub shear pads _ith no

restraint _ithin the hub, the S-mode inplane

shear and bending moments are reacted through

the elastomeric hub shear pad. The stiffness of

the shear pad has been tuned to accurately place

the first S-inplane frequency below 1 /rev (this

frequency for the current design is at approxi-

mately 0.6/rev) and well separated from the first

flap frequency at all operating conditions. The

damping in the shear pad elastomer provides a

high level of damping in the first S-inplane

motion. This, along with its large separation

from the 2/rev resonance condition ensures a low

level of blade dynamic loading for the 2/rev

Coriolis forces. In the C-mode inplane configura-

tion, the hollow construction of the hub and the

influence of the elastomeric shear pads allows the

flexbeam to bend within the hub. This ensures

that the bending moment loads are carried across

the hub within the flexbeam. Since the inplane

loads are not reacted by the shear pads in this

configuration, the first C-inplane frequency stays

well above 1/rev. The location of this frequency

and its damping can be optimized by proper choice

of flexbeam width, tip weight, pitch-flap coupling

and other parameters.

3. 2 Klexbeam Geometry

A rectangular flexbeam configuration was

chosen. Ho_ever, the span_ise distribution of

width and thickness were tailored for optimum

placement of fundamental 1-flap and 1-chord

frequencies as well as acceptable combined cor-

ner stresses. The "soft" hub mount of the flex-

beam and root-end kinematic pitcl_-lag coupling

ensured high damping of the rotor chord modes.

Hence no attempt was made to sandwich elasto-

merle material into the flexbeam design. The

chordwise stiffness was designed for adequate

separation of 1-chord and 1-flap frequencies.

The spanwise distribution of flexbeam width and

thickness has been configured for near uniform

spanwise distribution of combined corner Stresses

while maintaining a low structural torsional stiff-

ness. This is vitally important as can be seen in

Kig. 7, which shows a comparison of flapwise

bending stresses for different flexbeam configura-

tions for a blade flapping of {3 = 15 degrees.

Detailed calculations show that a flexbeam with a

uniform width and thickness is totally unacceptable

for fatigue loads at high for_ard Speeds.

3. 3 Tip Weight

The tip balance weight has been eliminated

for the CKTR. This results in a simpler tip

design _vithout a tip _veight attachment fitting.

Since the fundamental dynarnic effect is an

increased first C-mode chordwise frequency, the

removal of the tip weight is beneficial in separat-

ing the first flap and the first chord frequencies.

The spanwise balance _eight is located on the top

and bottom of the pitch case at its root end

(Station 10. 0). This location results in reduced

2_

tou I oULT

"_ _0_-.\

,o _\ /. um,o_ a.E:xsu_, ,b • ,L._m. :, • 0. P' m.,

01STANC[ FkO_l HUB SUPPORT {FRACTION OF FRff FL_XBEAM L_IGTHI

Kig. 7 Flapwise flexbeanl stress

{blade flap = 15 degrees}

feathering control loads due to reduced "tennis

racquet" effect.

3.4 Pitch Link Attachment

The pitch link is attached to the trailing edge

of the pitch case. For the design value of nega-

tive pitch-flap coupling {63 - -35 degrees), the

blade spanwise pitch horn attachment point is well

inboard, resulting in a small swashplate and a

compact design. In addition, the direction of the

pitch link load is the same as that of the rotor

thrust, thus reducing the flexbeam flap shear load.

Dynamically, because of the inboard attachment

of a trailing edge pitch link, the second flap

frequency is much higher than it would be for a

leading edge attachment. This is very important

in raising the second flap frequency above and

maintaining good separation from 3/rev. As

shown in Fig. 1, the pitch link is inclined radially

inwards from the s_ashplate to the pitch horn at

an angle of 70 degrees to the hub plane. This

induces kinematic pitch-flap-lag coupling to

improve the first inplane {C-mode) damping at

high collective pitch settings. The coupling

results in positive pltch-laK motion, i.e. , nose

down with blade lag motion. This is in general

agreement with the requirement for stiff-inplane

rotors.

3. 5 Chordwise Blade Balance

As in the existing AH-64 Inetal tail rotor the

chordwise c.g. of the CKTR blade has been

located at 35 percent chord to reduce the weight

of the blade and the "tennis racquet" loads on the

control system. Ballistic damage considerations,

ho_ew'r, require the rotor to be stable _ith a

failed pitch link. This condition is satisfied by

stabilizing the coupled pitch-flap mode with a

leading edge weight in the outboard portion of the

blade between 70 and 90 percent radius.

4. 0 Wind Tunnel Test Procedure

4. 1 General Description

The Composite Klexbeam Tail Rotor (CKTR)

was evaluated through extensive wind tunnel tests

to determine rotor performance, loads, and

dynamic characteristics in hover and in low and

high speed forward flight, and in sideslip condi-

tions that are representative of the production

AH-64 flight spectrum.

Testing was conducted in the Boeing Vertol

V/STOL wind tunnel located at Philadelphia,

Pennsylvania. The essential objectives of the

wind tunnel tests were:

1) Define dynamic and aeroelastic stability

characteristics of the CKTR over the sideslip

flight envelope of the AH-64.

Z) Define rotor loads, and blade load and

stress characteristics.

3) Define performance characteristics.

4) Define start/stop response

characteristics.

A fully instrumented blade pair assembly was

mounted on the Dynamic Rotor Test Stand (DRTS).

The DRTS assembly provided support, control,

and drive for the CFTR. A typical installation

with the rotor positioned for forward flight with

sideslip is shown in Fig. 8. Sideslip was simu-

lated by presetting the sting inclination, and

remotely controlling the DRTS pitch angle.

Twenty-six rotating gages were monitored. This

inchded flap, lag and torsion gages on the flex-

beam and the blade, pitch link, rotor hub, output

shaft, etc. Additional rotating and non-rotating

measurements include shaft torque balance thrust,

pitching and rolling moments, shaft angle, RPM

indicator, control system load, etc.

4. Z Control System and Rotor Support System

A close-up view of the drive and support sys-

tem is seen in Fig. 9. The test stand drive shaft

is coupled to the output drive shaft of the rotor

with adapting hardware. The "scissors" drive the

rotating s_vashplate from the output shaft.

The control system consists of the pitch link

attached to the pitch horn at one end and to the

rotating stxashplate at the other. The non-

rotating s_vashplate is mounted on two hydraulic

actuators {Fig. 9) spaced apart azimuthally by

180 degrees.

284

ORIGINAL PACE i_

DE POOR Q:'".".; :':

T h e s t a t i c m a s t i s moun ted on t h e D R T S with a n i n t e r f a c e h a r d w a r e ca l l ed the b a l a n c e a d a p t e r t h a t i s in t u r n s u p p o r t e d t o t h e t e s t s t a n d with a d y n a m i c b a l a n c e . T h e d y n a m i c b a l a n c e ( F i g . 9 ) i s s t r a i n - g a g e d t o m e a s u r e the C F T R t h r u s t , r o l l i n g and pi tching m o n ~ e n t s .

A d e s c r i p t i o n of t h e t e s t r o t o r c o m p o n e n t s i s p rov ided in R e f e r e n c e 9.

4. 3 Col l ec t ive and C y c l i c Exc i t a t ion

In p r e p a r a t i o n f o r wind t u n n r l t e s t s , p r o v i - s i o n w a s m a d e t o e x c i t e t h e r o t o r u s i n g co l l ec t ive a n d c y c l i c s h a k e r s . Th t , s e w e r e a v a i l a b l e to e x c i t e l owly d a m p e d f u n d a m e n t a l r o t o r m o d e s in o r d e r t o m e a s u r e t h e i r d a m p i n g c h a r a c t e r i s t i c s .

Cyc l i c m o d e s w e r e d r i v e n by a 300 Ibf, 0 - 2 0 0 Hz , s h a k e r moun ted on t h e s t i n g as shown in F i g . 8. T h e s h a k e r exc i t a t ion w a s app l i ed to t h e Dynamic R o t o r T e s t Stand ( D R T S ) below t h e s t a n d ba lance .

Co l l ec t ive exc i t a t ion w a s p r o v i d e d th rough t h e co l l ec t ive pi tch h y d r a u l i c d r i v e s y s t e m . T h e c o l l e c t i v e pi tch exc i t a t ion w a s u s e d with a n a m p l i t u d e of *O. 5 d e g r e e b l ade p i t ch change o v e r a f r e q u e n c y r a n g e 0 - 35 H z .

4. 4 T e s t P r e c a u t i o n s

F i g . 8 Co tnpos i t e f l exbeam t a i l r o t o r in thc. wind tunnel t e s t s e c t i o n

P r o c e d u r e s t h a t &'ere e s t a b l l s h e d t o e n s u r e t h e i n t e g r i t y of t h e C F T R t h r o u g h t h e c o m p l e t e t e S t enve lope included:

Non- ro ta t ing r a p t e s t s wer t . done a t t he s t a r t of e a c h d a y ' s t e s t i n g . i n t h e f l ap , l a g and t o r s i o n d e g r e e s of f r e e d o m wthre o b s e r v e d o n tht. s p e c t r u m a n a l y z e r . add i t ion to Trisual i n s p e c t i o n , t h i s t e s t p rov ided conf idence in the s t r u c t u r a l i n t e g r i t y of t he C F T R .

T h e r e s p o n s e of the blade

In

Se lec t ed r o t o r responsca gages m t ' r c ' con - t i nuous ly m o n i t o r e d f o r a l l t e s t cond i t ions on twe lve on - l ine m o n i t o r s and the s p e c t r u m a n a l y z e r . C e r t a i n c r i t i c a l g a g e s , in add i t ion to p e r f o r m a n c e p a r a m e t e r s , w e r e a l s o o b s e r v e d o n t h e on - l ine f l a tbed p l o t t e r s .

-

.4dditional t r s t p r o t e c t i o n w a s obscArved by i n t r o d u c i n g t h e c o l l e c t i v r pi tch d u m p capab i l i t y t h a t w a s dthsigned to a u t o m a t i c a l l y d r o p the c o l l e c t i v r pi tch t o a prcsviously t e s t e d s a f e l t ~ v t ~ l u h e n a n y one of se1rctc.d c r i t i c a l gagc r e s p o n s e e x c e cad ed a p r e s p t' c i f i ed va 1 u (1. a n a l y z e r s w e r e a l s o u s e d to con t inuous ly m o n i t o r tli e lion - h a r m on i c con t cn t of s el c c t c>d

r e s p o n s e s .

S pe c t r a 1

F i c . Q C F T R c!riv<. < i t r r I s u p p o r t 5 y" L < ~ t l l d b 5 c ' : :1 Ill).

This procedure for on-line data monitoring

and automatic colh-ctive pitch dump, safety of the

CFTR wind tunnel test was assured.

4.5 Test Stand Shake Test

Prior to mounting the CFTR on the Dynamic

Rotor Test Stand (DRTS), a shake test was con-

ducted to determine dynamic characteristics of

the test stand. The purpose of this investigation

was to:

1) Identify and isolate CFTR response

characteristics that were essentially the

influence of test stand dynamics.

2) I)eterrnine any distabilizing influence of

tee test stand on the rotor dynamics.

This was done by determining the test stand

frequencies, generalized masses, generalized

dampings, and mode shapes of all modes in the

frequency range 0 - 100 Hz. The hub modal data

was incorporated in a fully coupled CFTR/DRTS

aeroelastic stability analysis to w, rify that the

integrated systems are free from adverse

dynamic or aeroelastic coupling.

The influence of the test stand on the CFTR

modal characteristics were found not to be

significant.

4. 6 Data Reduction Eacility

Test data was processed for on-line or off-

line reduction and presentation. Off-line digitized

data was available in four formats.

1) Lo_ Speed Calculated Data presents

steady state static data of wind tunnel test con-

figuration. This data includes rotor advance

ratio, RPM, shaft antge collective pitch, C T,

Cp, velocity of wind tunnel, balance steadythrust pitching and rolling moments, velocity of

sound, etc.

2) High Speed Calculated Data essentially

calculates the steady and alternating values of the

different interaction equations {combined

stresses).

3) Stress Analysis Data presents the

steady and alternating values of _9 channels of

data being monitored for each test point.

4) Harmonic Am_iysis l)ata presents the

magnitude and phase of the first 10 harmonics of

all Z9 channels of data recorded.

Six on-line flatbed plotters _ere used to plot

any combination of dimensional or nondimen-

sional parameters in their final corrected forms.

Also available was on-line spectral analysis of

any selected data channel and corresponding

hard copies.

The wind tunnel control console offered

on-line monitoring of many ke.y control param-

eters. These were viewed in alphanumeric or

analog form on digital displays, oscilloscopes, or

oscillographs. A safety-of-flight monitor was

also provided. This data was continuously

recorded from a number of preselected data

channels whenew_r the rotor or tunnel was

activated. The parameters that triggered the

rotor blade pitch dump were monitored in analog

form on oscilloscopes.

5. 0 Kvaluation of Results

The test program determined the perfor-

mance, loads, and dynamic characteristics of

the CKTR for rotor speeds up to 1. 0 N R and air-

speeds up to 197 knots. The complete impressed

pitch range, as limited by test stand capabilities

or rotor structural requirements was investigated

in bow.r, low and high speed forward flight and

sideslip conditions. Static sideslip limits as

defined in the AH-64 System Specification (Refer-

ence 10) were investigated at airspeeds of 139,

164, and 197 knots. The stop/start characteris-

tics of the rotor in wind velocities up to 45 knots

were defined. The test explored the full steady

state sideslip envelope of the AH-64 as seen in

Fig. 10 where test points are superimposed on

the helicopters sideslip envelope.

120c-

_ 8o

oz_ 40

_-- -40"T(.9

12O

,I WIND TUNNEL TEST POINTS

o 40

_ ,_/,,,/ TRANSIENT LIMIT

I I I80 120 160

CALIBRATED AIRSPEED - KN

200

Fig. 10 AlI-64A sideslip envelope

For hover tests, the rotor speed was varied

from 0 to l. 0 N R (1403 RPM) in steps of 0. 2 N R

(4Z0 RPM). Collective pitch was varied over the

full range that was available at 0.8 N R, 0.9 N R

and 1.0 N R within the limits of the test stand

capability.

286

.0_. P_OOR QUALITY

Ot_t,_*-.,, * F_?.Cd. " "'

.OF £C,0!:, "'_ ....... '_

Fig. 11 presents a comparison of the CFTR

power versus thrust coefficient as measured in

the wind tunnel at zero wind tunnel speed.

Fig. 12 is the corresponding plot of rotor thrust

coefficient versus impressed blade pitch setting.

%x

mL_

ffou_

,=,

2

0.7

0.6

0.5

0.4

0.3

0.2

01

LEGEND

SYM RUN V KTS %RPM THETA S.S.

D 15 0 80 0

17 90

',) 18 100

o, 1004 0 02

Fig. Ii

&

%,

,it

,55'

0.02 004

THRUST COEFFICIENT, CT

Hover test, baseline blade

performance data

0.06

Forward flight tests were conducted for the

conditions shown in Fig. 10. Sideslip angles at

V = 138 knots, 164 knots and 197 knots were

essentially restricted to the steady sideslip

limits. Attempts to test at higher left and

right sideslip angles resulted in autorotation of

the rotor for zero collective pitch. This, of

course, is a test stand limitation and will not be

encountered in actual flight.

Typical spanwise distribution of flexbeatn

and blade loads at V = 164 knots and _3SS - +6

degrees is shown in Figs. 13 through 18. Pitch

case loads (station 4. Z to Z5. 0 inches) are not

shown in these figures since it was not instru-

mented. Flexbeam loads for various pitch

angles are shown between station 6. 2 inches and

25. 0 inches and the blade loads between station

25. 0 inches and 56.0 inches. The pitch case,

Q

<k-

JL3Z<

K

<

m

3o

2o

10

.fo --

2O

004

LEGEND

SYM RUN V KTS %RPM

15 0 80

1 7 90

_' 18 100

THETA SS.

0

Z1

-0 02 0

G

,3

i I0.02 0 04

THRUST COEFFICIENT, CT

Fig. 12 Hover test, baseline blade

performance data

0 06

500

400

300 --

- -j_

I-

E

_ -2o

.23o

3_c

Fig.

LEGEND

SYM v KTS %RPM THETA SS

C ',4 :t"7

2

I I I I Ilo 2c 30 4C _,C 6c

BLADE STATION INCHES

13. CFTR _ind tunnel test - mean flapmoment distribution

2 8 7

flexbeam, and blade junction is at station 25. O.

These stations are important in understanding

the discontinuities and inflections in the bending

nloment plots.

The steady loads between the pitch case, flex-

bcazn and blade should balance at the junction,

station 25. 0. Ho_ever, because of phase differ-

ences between the loads in the pitch case, flex-

beam, and blade, the plots of the oscillatory loads

do riot necessarily add up at the junction.

400

io

×

z

I

w

O

300

200

LEGEND

SYM V-KTS %RPM THETA S.S.

e 164 100 O 6

-_B--- 2

ZN 8

100

0 10 20 30 40 50


Fig. 14 CFTR wind tunnel test - alternating

flap moment distribution

6o

\

\

\\

L_GE_ID

SYM V KTS _R PP,A THfTA SS

C 2

O &

I I I i I1j ?c, 30 40 _,_

BLA_;[ ST,_TIO% INt HI!,

Fi_. 1 5 CFTR _ind tunnel test - mean

chord nlomcnt distribution

o

×

J

z

z

o

L;

<

T

9

5z_

p-zLU

o

gm

288

900

800

700

600

LEGEND

SYM V-KTS %RPM THETA $.S.

O 164 100 0 6

O 2

© 6

%

\bOO -- \

\\

400 -- X

< _ \\\

ioo

o I I i I i I(_ ILl 20 30 40 50

V,LAI)E STATION INCHES

Fig. 16 CKTR wind tunnel test - alternating

chord moment distribution

Joo

200

100

-100

-200

JOO

LEGEND

SYM v KTS %RPM THETA S,S.

0 164 100 0 6

E ?

O 6

S

I I I I I10 20 30 40 50


Fig. 17 CFTR wind tunnel test - mean

torsion moment distribution

6o

6o

OF POOR QUALITy

ORIGINAL PAGE Ig

OF POOR QUALITY

160

140

120

×

=,--i_ lO0

r

O 8o

g

6o

h<

4O

RUN

43

LEGEND

SYM T.P V KTS %RPM THETA S.$.

O 2 164 10U 0 6

FJ 3 p

O 5 6

I I l I I1 L} 20 30 10 50 50

_ILA[)E SrATIO'. t%CH! S

Fig. 18 CFTR _ind tunnel test - alternating

torsion moment distribution

Steady and alternating flapwise bending

moments are shown in Figs. 13 and 14. Both

show a steep drop in flexbeam bending moment

from the edge of the hub to approximately station

10. 0 inches. As per design, the flexbeam flap

bending moment tapers to practically zero between

station 20.0 inches and 25. 0 inches. The jump

discontinuity in the bending moment between the

flexbeam and blade at station 25. 0 is the b_mding

moment in the pitch case. The flapwise

bending moment in the pitch case would reduce to

zero at the pitch link/pitch horn attachment.

Similarly, the bending moment distribution is

drawn such ti:at the value at the blade tip {station

56. 0 inches) is zero Chordwise bending

moments are seen in Figs. 15 and 16. The dis-

continuity at station 25. 0 inches reflects the

chordwise loads in the pitch case. The component

of pitch link compression load in the chordwise

direction produces this bending moment. The

chordwise load in the pitch case is essentially the

result of thu pitch link inclination. Unlike the

flap bendin_ moment distribution, the chord_ise

moment in the flexbeam has a more gradual dis-

tribution. The torsion bending moments are

shown in Figs. 17 and 18. The steady flexbeam

torsion load is due to the steady wind-up of the

flexbearn. Measured flexbeam torsional load for

03/4 = 8 degrees is approximately zero since the'

flexbeam is unt_isted at this pitch setting. The

difference between the blade and flexbeam torsion

bending moment at station 25. 0 inches is the tor-

sion load in the pitch case reacted by the pitch

link. Fig. 17 also sho_s the relative magnitude of

the flexbeam torsion load to the pitch link load.

Alternating torsion load in the flexbeam is a

result of flexbeam feathering with blade flapping

with the root-end pitch flap (63} coupling.

5. 1 Dynamic Results

As discussed in Section 4. 3, collective and

cyclic shakers were available to excite lowly

damped fundamental rotor modes in order to

measure their damping characteristics.

The collective pitch excitation had an ampli-

tude of +0. 5 degrees blade pitch change over a

frequency range of 0 - 35 Hz. The cyclic excita-

tion was input as non-rotating test stand force

with the 300 lbf shaker. Shaker forces of 50 lbf

and 100 lbf were used from 0 - 70 Hz.

Accordingly, collective and cyclic excitation

were attempted to excite the rotor modes at each

point in hover in the test envelope. However,

after many attempts it x_as determined that the

rotor fundamental modes were heavily damped

and, hence, could not be excited with either of the

two shakers. It was decided at this point that

envelope expansion of CFTR wind tunnel test

_ould be based on the magnitude of non-harmonic

flap, lag or torsion response as seen on the

on-line spectrum analyzer.

Dynamic analysis research tool (DART)

analysis program was used to define the CFTR

dynamic and aeroelastic characteristics and blade

loads of the CKTR. This program is describedin Reference 11.

Two basic types of analysis w'ere used to sub-

stantiate the dynamic and aeroelastic character-

istics of the CFTR. First, an eigenvalue analysis

was used for configurations in how,r to establish

freedom from aeroelastic instability throughout

the complete blade pitch and rotor speed ranges

of the CFTR. This also established the blade

modal characteristics. Second, forward flight

stability" _vas established by trimming the rotor at

OE poOR QUALITY.

different points of tile flight envelope. Since the

analysis included nonlinear structural couplings

and aerodynamics (including dynamic stall), rotor

trim without nonharmonic response indicated

positive stability margins.

The resonance diagrams generated by DART

for reactionless, cyclic and colh-ctiw, boundary

conditions are shown in Figs. 19, 20 and 21,

respectively. Test frequencies obtained at zero

and operating RPM are superimposed on the

resonance diagrams.

Tabulated results of the non-rotating rap tests

are shown in Table i. The fundamental l-flap,

2-flap, l-chord and l-torsion modes show good

correl.ation with analytical data. Spectral plots

of non-rotating rap tests for flexbeam chord and

flap gages are shown in Figs. 22 and 23,

respectively.

Results of cumulative spectrum plots for

different for_ard flight tests are shown in

Table 2. Spectral plots for one flexbeam chord

gage for V = 1 _9 knots and 197 knots are sho_n

in Figs. 24 and 25, respectiwqy.

_om_S,EEe@:Zo,

Fig. 20 CFTR resonance diagram - cyclic

modes, 03/4 = 0

: -

.o_o.s,,_D,u._

d bM

L J

Fig. lq CFTR resonance diagram - reactionless

modes

Fig. gl CFTR resonance diagram - collective

modes, @3/4 = 0

2go

DF :-C:(3_:QUALITY

Table 1. Nonrotating Modal Frequencies - Correlation

of Test Results with Analysis

Frequency - Hz /Rev)

Configuration Mode Analysis Test

React_onless 1-Flap 3. 5 (0. 15) 4.4 (0. 19}

Boundary 1-Chord 16.4 (0.7) 18.2 (0. 78)Condition

Cyclic 1-Chord 30.4 (1.3) 32.3 (1.38)

Boundary 1-Torsion 51.4 (2.2) 53.8 (2.3)

Condition i-Flap 69.0 (2.95) 70.2,(3.0,

66.4 2.84)

Collective 1-Torsion 40. 9 (1. 75) 40.0 (1.71)

Boundary l-Flap 57. 3 (2.45) 58. 0 (2.48)Condition

Table 2. Inplane Modal Frequencies for Various Test Conditions

Test Condition

Figure Collective

No. V (KTS) {3s s (Deg.) Pitch

Flexbeam Chord Gage

Resonant Frequencies

Hz ( /Rev)

25 -90 Sweep 8.4 (0. 36/Rev); 17.8 (0.76 Rev);

33.0 (1.41 Rev)

24 139 +15 Sweep 6.8 (0.29/Rev); 16.8 (0.7i/Rev);

29.0 (1.24/Rev); 70.0 (3/Rev)

25 197 -8 Sweep 7.5 (0.32/Rev); 15.2 (0. 65/Rev);

Z9.7 (l.27/Rev); 70.0 (3/Rev)

197 - g Sweep 7.5 (0. 32/Rev); 15.7 (0. 67/Rev);

29.5 (1.26/Rev); 70.0 (3/Rev)

0-164

Sweep

0 0 7. 7 (0. 33/Rev); 16.8 (0. 7g/Rev);

30. 5 (1. 30/Rev); 70.0 (3/Rev)

291

--53 8 Hz {E.3/REVJ

58.0 Hz (2.5/REV}

I F-%0 Hz (LINE FREQUENCY)

F 4 5 Hz (0.19/REv) H

i I £,oH,_3e,,.,tv,

5798fi Hz 100

Fig. 22 Frequency spectrum of flap response -

nonrotating rap test

Hz (0 78/REV)F-,83| F-3_3 .,. 38,.EVl 1

_L / / _ 40.0Hz(1.71/REV'

t r ;Fig. Z3 Frequency spectrum of chord response -

nonrotating rap test

23.4 Hz (1,'REV)

16.8 Hz /

t0.72/RE V)ji (3.0/REV)

,_2oo.. ,1 ;6 8 Hz I1 24/REV) 468 Hz ! i (4/RE_J /

(0 29/REVI I 1 (2/REV) : , "' 1-

3900_

Fi_ 24 Frequency spectrum of flexbeam chord response -

V = 139 knots, [3ss = +15 degrees, 0 sweep,

f2 = 1403 rpm

f 234 Hz {(1/REV}! ,%"_,]1_,.,.00,,._ li 20_., ,68.,

J_Io.32/REVlt , 1/'\ 'I | _ 4-

__ ......_ _ _b--+ _ ] '__'_'_ L_3000 _ Hz 100

FiA. Z5 Frequency spectrum of fle×beam chord response -

V = 197 knots, _ss = -8 degrees, @ sweep,

S2 = 1403 rpm

292

O":;GTE_AL PAC-E [_

OF POOR QUALITy

OIP._C-.'}....,_ ', :f-::.-?3 _5O1 ..........

5. 1. 1 Reactionless Boundary Condition

The reactionless boundary condition corre-

sponds to an isolated rotor. The reactionless

modes resonance diagram for the collective pitch

extremes of -14 degrees and +27 degrees is

shown in Fig. 19. In the reactionless or

"scissors" (S-mode} inplane boundary condition,

the steady and 2/rev inplane shear and bending

moments are reacted through the elastomeric hub

shear pads. The stiffness and damping of the

shear pads provide the hub restraint for blade

chordwise motion. The first chord frequency is

primarily dependent on the stiffness and span-

wise offset of the hub shear pad. Its frequency

is located at approximately 0. 6/rev which pro-

vides good separation from the first flap fre-

quency and 2/rev Coriolis excitation. The first

flap frequency is governed by the effective hinge

offset (approximately 10 inches} and the value of

kinematic pitch-flap coupling. The first flap is

generally highly damped. The high damping of

the first chord mode is a reflection of hub shear

pad damping characteristics. This is evidenced

by the results of shake tests using the collective

and fixed system shakers. Since the hub shear

pads do not feather with pitch change, the first

chord frequency and damping remain essentially

unchanged with change in blade collective pitch.

The first flap frequency and damping are gen-

erally unchanged with collective pitch.

The higher modes have been shown analyti-

cally (Reference ll) to be well damped with

minimal change with collective pitch.

The coupled mode shapes corresponding to

the fundamental modes are shown in Figs. 26 and

27. The first chord mode, ICig. 26, shows very

little coupling with the flap and torsion motion of

the blade. The elastic deflection in the chord-

wise direction is essentially in the hub shear pad

with the blade moving as a rigid body. The first

flap mode, Fig. 27, shows the coupling between

the blade flap and torsion motion (pitch/flap

coupling}.

In contrast to conventional rotors, the first

torsion mode reflects feathering motion about the

pitch link/pitch horn attachment. The shear

stiffness of the snubber in flap and chord and the

chordwise stiffness of flexbeam between station

15. 0 inches and 25.0 inches, in addition to the

control system stiffness, have significant influ-

ence on the frequency of this mode. This is

determined from the strain energy data corre-

sponding to the first torsion mode.

UP, LAG, NOSE UP

3.0

2.0

1.0

ox

,,_ 0.0iI:

1.0

2.0

3.0

1 CHORD MODE

FREQUENCY (CYC/REV) = 0.612

-- CRITICAL DAMPING RATIO = 0.0_

C.O.DToRS,ON./

I FLAP

-- IFLEXBEAM AND ' BLADEPITCH CASE

l I I I I10 20 30 40 50 60

BLADE STATION IN.

Fig. 26 Reactionless B.C., mode

shape plots -- l-chord mode

UP, LAG, NOSE UP

3.0

2.0

1.0

o

x

< 0.0

1,0

2.0

1 CHORD MODE


CRITICAL DAMPING RATIO_"_'_ FLAP

D

FLEXB_

- PJCHCASE/ II

FLEXBEAM AND"_--_ _" BLADEPITCH CASE

"_""_ TO RSIO N

30 I I I I I0 10 20 30 40 50

BLADE STATION - IN.

6O

Fig. 27 Reactionless B.C., mode shape

plots -- l-flap mode

5. 1.2 Cyclic Boundary Condition

In the cyclic or C-mode boundary condition,

the 1/rev inplane bending moments are contained

within the flexbeam in the carry-through hub con-

struction and are not reacted through the hub

shear pads and the hub. The hub support flexibil-

ity is n_odeled. The coupling between the hub

293

>

z

3

motion and blade" feathering due to swashplate

motion is included. The kinematic flap-lag-

torsion coupling due to pitch link/pitch horn

spanwise and chordwise location and pitch link

inclination is also included in the analysis.

The regressing frequencies for zero collec-

tive pitch are shown in Fig. 20. The first chord

frequency, which reflects the stiffness of the

flexbeam and the inertia of the blade, is well

separated from the first flap frequency and from

I/rev resonance.

Fig. 28 shows the influence of collective

pitch on blade frequencies. The first flap fre-

quency remains practically unchanged with collec-

tive pitch. The pitch orientation of the flexbeam

with respect to the blade chord ensures minimal

variation of the first chord frequency over the

collective pitch range of the rotor. The first

torsion mode shows a drop in frequency with

collective pitch thus further separating it from

3/rev. As expected, the second flap frequency

increases and the second chord frequency

decreases with changes in collective pitch from

zero.

10

LEGENU

MEASURED wqND _UNNL L

9 -- _EST .!}_! 8_ ) 7( _FRECUEN(:I[S • ."

• " 6_Z# a._4" .'° ."-- , •------_ _2j (_ ."

I

5_2/ ."

4_2

..°. 3

2NI_ FL_-' ". • " ." .'" ..''

C_t_}HD

...." .' .. .." ,.""

_).? o4 cl_ oa i u I ? _ 4

Fig. 28 CFTR resonance diagram; cyclic

modes @3/4 = -14 and fi7 degrees

Figs. Z9 through 31 show the fundamental

coupled mode shapes for the cyclic boundary con-

dition. The first flap mode, Fig. 29, shows the

pitch/flap coupling for cyclic boundary condition.

The first chord mode shows the amount of kine-

matic pitch/lag coupling induced by the inclined

pitch link. The first torsion mode, Fig. 31,

shows the extent of flap coupling.

UR LAG, NOSE UP

3.0

2.0

1.0

o

x

0.0cE

1.0

2.0

3,0

1 FLAP MODE (REG.)

-

-- _ FLAP

jC.ORD!

I _" TORSION- I

FLEXBEAM AND I

PITCH CASE _ BLADE

I I I I I10 20 30 40 50 60

BLADE STATION - IN.

Fig. 29 Cyclic B.C., mode shape

plots -- 1-flap mode

UP, LAG

3.0

2.0

1.0

o

x

0,0

1.0

2,0

3.0

NOSE UP

1 CHORD MODE (REG.)

FcRE,T_; EA_CDY(C _,C'2 ERdAT, O 11;40825

ORD

_FLAP

\II TORSIONII

Fl" EXBEAM AND IPITCH CASE BLADE

I I I I I I

10 20 30 40 50 60

BLADE STATION IN


plots -- 1-chord mode

29_

UP, LAG,

3.0

2.0

1.0

O

x

O.O

1.0

2.0

3.0

NOSE UP

1 TORSION MODE (REG.)


CRITICAL DAMPING RATIO = 0.07

-- S FLAP

FLEXBEAM AND IPITCH CASE _ BLADE

I I I I I10 20 30 40 50

BLADE STATION IN,


plots -- l-torsion mode

6O

This was achieved through placement of the

reactionless 1-chord frequency below l/rev.

Comparisons of harmonic loads between the CFTR

and a similar size rotor (Reference 1Z) based on

test data are seen in Figs. 32 through 35.

Figs. 32 and 33 are flight test loads of the YUH-

60A tail rotor. Figs. 34 and 35 are wind tunnel

test loads for the CFTR. This comparison is a

study of the relative magnitudes of the harmonic

loads for geometrically similar rotors with differ-

ent dynamic characteristics. Absolute magnitudes

of the loads should not be compared. The span-

wise distribution and relative harmonic content of

flapwise ftexbeam loads are similar between the

two rotors (Figs. 32 and 34). However harmonic

contents of chordwise loads between the two rotors

are quite different. In Fig. 33 (stiff inplane rotor),

chordwise 2/rev loads are higher than the 1/rev

loads. The CFTR (Fig. 35, soft inplane rotor)

chordwise Z/rev loads are an order of magnitude

lower than the 1/rev loads. This trend has been

found for all test conditions.

5. 1.3 Collective Boundary Condition

The difference between the collective and

reactionless boundary conditions are in the model

for the control system and drive system. The

drive system torsional flexibility is represented

by its flexibility in the blade inplane structural

model at the hub. The control system stiffness

is reflected by the structure from the tail rotor

actuators to the pitch horn. The effective mass

of the swashplate assembly has a significant

influence on the first torsion frequency.

The resonance diagram for the collective

boundary condition is shown in Fig. il for zero

collective pitch. The predicted first chord modal

frequency, which is essentially the drive system

torsion mode, is omitted in the plot. This is

because the frequency and damping of the first

chord mode is more accurately predicted in the

stability analysis of the tail rotor drive system

rather than from the rotor model. The drop in

the frequency of the first torsion mode {from

those of the reactionless boundary condition) is a

reflect'ion of the reduction of control system stiff-

ness and the inclusion of swashplate assembly

inertia for the collective boundary condition. The

second chord frequency is also reduced as a result

of tors-ionat flexibility of the drive system.

Experimentally determined 1-chord frequency is

included for comparison.

• 5-2 Harmonic Loads

As discussed in Section 3. 0, the CFTR was

designed for low chordwise 2/rev Coriolis load.

5

4

%x

3

F-

_2o

00

.-SNUBBER / £ B,R

I

i _ 1ST HARMONIC

-- _ 3RD HARMONIC

\10 20 30 40 50 60 70

BLADE SPAN _ r/R _ PERCENT

Fig. 32. YUIt-60A tail rotor blade harmonic

analysis flatwise (V = 143 KTS)

295

£

-J 2

I

O

O

>

00

_%,% ,,_ SNUBBER F CL RIB

_ _ ! 2ND HARMONI i

1 ST HAR

10 20 30 40 50 60 70

BLADE SPAN _ r/R _ PERCENT

Fig. 33 YUH-60A tail rotor blade harmonic

analysis edgewise (V = 143 KTS)

40o

x

.J

3.o

w

o

20m

m

o

w

g

V = 139 KT

_ss = +15DEG

03/4 = + 16 DEG

CT = .0089

SYM HARMONIC

© 1P

[:3 2P

O 3P

Z_ 4P

EDGE

OF

HUB

ii

0

"_ PITCH CASE _..p--BLADE

" "--'-'"rh--P -- : -r ''''---_' _10 20 30

R (INS)

Fig. 35 CFTR - ftexbeam harmonic loads

2400

2000Z

1600

1700

Osoo

40O

SYM liARMONIC

"V T39 KT 1P

_S$ • 15 DE(32P

03/4 • 16 DEG\ 3P

C T 0089 _'- 4P

EDGE

OF

HUB

PITCH CASE -_11---I_

FLEXBEAM _ _ BLADE

10 20 30 35

R (INS)

Fig. 34 CFTR - flexbeam harmonic loads

6.0 Concluding Remarks

As discussed in the preceding sections, the

HHI Composite Flexbeam Tail Rotor has a

dynamically unique design. This rotor has beendemonstrated, through wind tunnel tests, over

the full sideslip envelope of the AH-64, Advanced

Attack Helicopter. The wind tunnel tests havevalidated that the CFTR:

1) Is aeroelastically stable throughout the

complete collective pitch range and up to opera-

tional rotor speed of 1403 RPM.

2) Is aeroelastically stable for fo:ward

flight speeds up to 197 knots and sideslip flight

representative of the AH-64 flight envelope.

3) Has excellent dynamic characteristics

at all pitch angles, rotor speeds and test

conditions.

4) Exhibits low flexbeam flapwise and

chordwise steady and alternating stresses.

Loads were well below endurance limit for all

conditions tested in the wind tunnel.

5) Does not require a complicated flex-

beam cross-section design with elastomeric

material sandwiched in the flexbeam to provide

damping.

These excellent characteristics have been

achieved through judicious choice of design

innovations which are the result of industry

experience with bearingless rotors. Some of

these innovations are discussed below:

1} In order to avoid stability problem char-

acteristics of bearingless tail rotors, the first

inplane reactionless (S-mode) frequency was

tuned below 1/rev while maintaining the first

inplane cyclic (C-mode) frequency above 1/rev.

Both frequencies are well separated from the

first flap frequency. This was accomplished

through the design of the chordwise stiffness of

the flexbeam, and by elastomerically mounting

the flexbeam to the hub.

Z) By allowing the flexbeam to freely flex

within the hub, the load transfer to the hub is

minimized. The 1/rev chordwise load is main-

tained within the flexbeam and not transferred to

the hub. The 2/rev chordwise loads are trans-

ferred to the hub after significant attenuation due

296

D]_ P.OOR QUALITY

0,_ PCGK QJALITZ

to hub shear pad damping and separation of the

first chord reactionless frequency from 2/rev.

3) The trailing edge pitch link attachment

was found to be advantageous over a leading edge

configuration (for a bearingless rotor of the

"pusher" type}.

a) For the required kinematic pitch-

flap coupling of -35 degrees, the trailing edge

pitch link attachment permits a smaller swash-

plate and a compact control system design.

b) The trailing edge pitch link attach-

ment raises the second flap frequency, thus pro-

viding good separation from 3/rev.

c) The nominal pitch link load (com-

pression) for a trailing edge pitch link attachment

is in the same direction as the rotor thrust, thus

reducing considerably the flap shear load in the

flexbeanq, inboard of the pitch shear support.

4) The inclination of the pitch link intro-

duces positive pitch-lag coupling {nose down _ith

blade lag). This coupling adds damping to the

first chord cyclic mode through pitch coupling,

especially at higt_ collective pitch settings.

5) The relative pitch orientation of the flex-

beam chord with respect to the blade chord causes

the cyclic first chord frequency to first increase

and then decrease through the collective pitch

range of the rotor. This ensures minimum

decrease of the cyclic first chord frequency and

prevents coalescence with the first flap frequency.

6) The above means of introducing damping

and of preventing dynamic instabilities involving

the lowly damped I-chord mode, eliminates the

need for introducing structural damping through

elastomeric inserts in the flexbeam.

7) The leading edge balance weight between

station 39 and 51 was introduced to move the blade

dynamic center of gravity forward and eliminate

blade flutter due to structural failure of the

feathering control system.

8) The blade spanwise balance weight is

located at station 9.7 (on top and bottom of pitch

case) rather than at blade tip. Elimination of

tip balance weight increases the cyclic first chord

frequency and avoids coalescence with the first

flap frequency. The balance weights on the top

and bottom surfaces of the pitch case act as

"Chinese" weights, thus reducing feathering

control loads.

I.

7.0 References

Edwards, W.T. and Miao, W., "Bearingless

Tail Rotor Loads and Stability", Boeing-

Vertol Company, prepared for Applied

Technology Laboratory, Research and

Technology Laboratories (AVRADCOM),

Fort r2ustis, VA 23604, USAAMRDL

TR 76-16, November 1977.

Z. Fenaughty, R.R. and Noehren, W.L. ,

"Composite Bearingless Tail Rotor for

UTTAS", Journal of the American Helicopter

Society, July 1977.

3. Cook, C.V., A Review of Tail Rotor Design

and Performance, Vertica, Vol. 2,

pp. 163-181, 1979.

4. Hughes, C.W., "Design and Testing of a

New Generation Tail Rotor _', Bell Helicopter

Textron, presented to the AHS Technical

Council for consideration of the Robert L.

Lichten Award, February 1978.

5. Maloney, P.F. and Porterfield, J.D.,

Elastic Pitch Beam Tail Rotor, Kaman

Aerospace Corporation, USAAMRDL

TR 76-35, U.S. Army Air Mobillty Research

and Development Laboratory, Fort Eustis,

VA 23604, December 1976.

6. Gaffey, T.M. , "The Effect of Positive

Pitch-Flap Coupling (Negative 63) on Rotor

Blade Motion Stability and Flapping",

Journal of American Helicopter Society,

April 1969.

7. Ormiston, R.A. and Hodges, D.H.,

"Linear Flap-Lag Dynamics of Hingeless

Rotor Blades in Hover", Journal of

American ttelicopter Society 17 (2),

April 1972.

8. Head, R.E. and Banerjee, D., "Helicopter

Tail Rotor of the Elastomerically-Mounted

Composite Flexbeam Type", Patent

No. 4,381,902, May 1983.

9. Banerjee, O., "Composite Flexbeam Tail

Rotor for the AH-64 Advanced Attack Heli-

copter Wind Tunnel Test Report", Hughes

Helicopters, Inc., Report No. 150-V-2003,

HHI 82-362, December 1982.

10. Systems Specification for Advanced Attack

Helicopter, AMC-SS-AAH-H10000A.

11. Banerjee, D., "Composite Flexbeam Tail

Rotor for the YAH-64 Advanced Attack

Helicopter, Aeroelasticity and Rotor Blade

Loads Report", Hughes Helicopters, Inc.,

Report No. 150-V-2001, HHI 82-186,

June 1982.

12. YUH-60A - Structural Load Survey,

Sikorsky Aircraft Report No. SER-70406.

13. Huber, H., Frommlet, H., and Buchs, W.,

"Development of a Bearingless Helicopter

Tail Rotor", Sixth European Rotorcraft

and Pouered Lift Aircraft Forum, Paper

No. 16, September 16-18, 1980, Bristol,

England.

297

Documents

N88-27162 - NASA